The present invention relates to a method for producing a central wing box.
As illustrated in FIG. 1, the structure of an aircraft comprises two sub-assemblies, on the one hand a fuselage 10, and on the other hand a wing 12, that are linked by means of a box structure 14 known as a central wing box.
As illustrated in FIG. 2, the central wing box 14 comprises on the one hand two panels, an upper panel 16 and a lower panel 18, and on the other hand, a minimum of two spars, a front spar 20 and a rear spar 22.
In a known manner, these four elements are produced separately and then connected to one another, using intermediate parts 24 such as angle members, as illustrated in FIGS. 2, 3A and 3B and/or via lightweight extensions 26 at the level of the panels or spars of the edge flange type, as illustrated in FIGS. 3B and 3C.
In all cases it is necessary to provide at least two joints per spar, i.e. at least four joints for the central wing box 14.
In the case of metal parts, each joint necessitates a phase of positioning of the two elements to be assembled, a pre-assembly phase, for example by pinning, a drilling/boring phase, a phase of cleaning off swarf, and a phase of riveting several rows for fixing.
These operations are long and have a significant impact on the cost of the box.
In the case of panels and spars made of composite material, the production method comprises the same steps as for metal elements, but the drilling phase is even longer and more difficult because of the risks of flaking.
According to another problem linked to the composite material, the gap between the parts to be assembled should be less than 3/10 mm in order to obtain a contact between the parts enabling them to be assembled without residual bending stresses. This stress necessitates perfect control of the methods of manufacture of the parts to be assembled, in particular in the region of contact surfaces.
However, it is quite often necessary to provide an additional step consisting of interposing a wedging resin between the two parts to be assembled in order to be within the contact tolerance.
This operation is long since after the resin has been deposited the parts must be assembled temporarily in order to calibrate the thickness of the resin, then disassembled for drying of the resin. The final assembly is performed only after the drying of the resin.
According to another point, at the edges of the parts to be assembled, assembly by riveting generates substantial local stresses that necessitate extra thicknesses. In the case of composite material, these extra thicknesses are more substantial since the orientations of the fiber reinforcements must be optimal. Of course, these extra thicknesses increase the payload.